1. Field of the Invention
The instant invention relates to a sealing device for the turbine blades of a turbojet engine, specifically such sealing devices which are adjustable to maintain a specific clearance between the sealing structure and the turbine blade tips during all operating modes of the tubrojet operation.
2. Brief Description of the Prior Art
It is important to minimize the clearance between the turbine blade tips and a sealing device in a turbojet engine in order to maximize the efficiency, to maximize the thrust, and to obtain a satisfactory surge margin of the turbojet engine due to leaks in the clearance between the rotating and stationary parts of the engine.
In order to reduce the leakage between the turbine blade tips and the surrounding structure, it is necessary to reduce the clearance between the blade tips and the sealing device to a minimum dimension and to maintain this dimension in both stable and transitory engine operating modes. The sealing device must remain concentric with the axis of rotation of the turbojet engine, and must expand and contract in a radial direction to compensate for the expansion and contraction of the turbine blades. The blades will undergo expansion during engine acceleration due to the increase in centrifugal forces and due to the increases in operating temperatures. Conversely, the turbine blades will contract during periods of engine deceleration or stabilized low power operating modes.
It has been extremely difficult to design a sealing system that surrounds the turbine blade tips and maintains a predetermined, minimum clearance during all stages of the turbojet engine operations. In addition to compensating for the expansion and contraction of the turbine blade tips, the sealing device must also take into consideration the potential action of inertia forces acting on the aircraft engines (load factors in the Z or Y direction) and deformations due to changing thermal characteristics. Additionally, the sealing device must retain its circular shape and cannot assume any degree of ovalness without incurring the risk of contact between the sealing device and the blade tips. Such contact would, at best, cause increases in the leakage between the blade tips and the sealing device and could possibly cause severe damage to the turbine blade structure.
The prior art devices have attempted to achieve these objectives by constructing a very rigid and heavy, or a very complex sealing system. Both systems have obvious drawbacks in regard to their use on aircraft engines: the first serving to increase the weight of the aircraft; while the second decreases the reliability of the turbojet engine.
The prior art also includes systems utilizing an abradable sealing surface which is worn away by the action of the turbine blades to minimize the clearance between them. However, these systems have not alleviated the leakage problems since, during expansion of the turbine blade tips, they abrade away the sealing surface and, when the operating conditions such that the turbine blades contract, a large clearance between the blade tips and the sealing device is present. An obvious way of avoiding this problem is to design the sealing device to accommodate the maximum diameter of the turbine blades. However, this introduces excessive leakage during those periods of operation when the turbine blades are not at their maximum diameter.
Although it is known, as described in French application No. 81.20719, filed Nov. 5, 1981, to center the casing supporting the sealing device with respect to the axis of the turbojet engine and provide it with sufficient inertia so that its deformation is essentially negligible, such devices cannot maintain a positive, but very small clearance between the sealing device and the turbine blade tips during both transitory and stabilized operating modes of the turbojet engine.
The prior art has also attempted to adjust the diameter of the sealing device in order to accommodate the expansion and contraction of the turbine blade tips by directing air taken from one or more stages of the turbojet engine compressor onto the sealing device to thereby cause its thermal expansion or contraction in a radial direction. The air is first directed into a distributor which, in turn, distributes the air in a homogeneous manner about the periphery of the sealing device. However, the quantity of air that is necessary to achieve the expansion or contraction of the sealing device in order to accommodate for both the centrifugal expansion of the turbine wheel and turbine blades (which occurs in a few seconds) and the subsequent thermal expansion of the turbine wheel (which takes place over several minutes) is usually excessive and results in the decreased efficiency of the turbojet engine compressor. A typical showing of such a system appears in French Pat. No. 2,467,292.
Although such air distributors can obviously be designed, as the prior art has indicated, they are extremely complex and, consequently, rather unreliable. Needless to say, a failure of such distributor would result in severe damage to the turbine blade or the sealing device.
As typlified in French Pat. Nos. 2,450,344 and 2,450,345, it is known to attempt to solve the problems noted above by making an inner part of the sealing device expand or contract to accommodate for the rapid centrifugal expansion of the turbine wheel and the turbine blade during acceleration and a second part which accommodates for the thermal expansion of the turbine wheel. However, such devices have been applied only to relatively low power turbojet engines having reverse flow combustion chambers. Although, in theory, such a system could be applied to the usual direct flow chambers of high power turbojet engines, they would be unduly complicated and inherently unreliable.
It is also known to utilize an elastic sleeve disposed about the turbine blades which is capable of deformation when exposed to stress. However, the elasticity of the sleeve presents the risk of introducing damage due to the lack of concentricity with the turbine wheel rotational axis, and due to the oval shape under the effect of load factors encountered in flight. It should be further noted that with the considerable hyperstatic forces generated by the supports in a segmented annulus, such as that shown in French Pat. No. 2,450,345, the slightest heterogeneity in temperature or inertia of the annular structure in the peripheral direction, will cause substantial deformations of the segmented ring. Such deformations will cause either lack of concentricity or result in the ovalization of the sealing structure, two factors, the maintenance of which are absolutely necessary to prevent excessive clearances between the turbine blade and sealing device.